The field of the disclosure relates generally to gas temperature measurement, and more specifically, to methods and a system for measuring gas temperature in harsh environments based on radiation thermometry using thin filaments.
At least some known turbomachines, such as gas turbine engines, include a plurality of rotating turbine blades or buckets that channel high-temperature fluids, i.e., combustion gases, through the gas turbine engines. Known turbine buckets are typically coupled to a wheel portion of a rotor within the gas turbine engine and cooperate with the rotor to form a turbine section. The turbine buckets are typically spaced circumferentially in a row extending about the rotor. Moreover, known turbine buckets are arranged in axially-spaced rows that are separated by a plurality of stationary nozzle segments that channel the fluid flowing through the engine towards each subsequent row of rotating buckets. Each row of nozzle segments, in conjunction with an associated row of turbine buckets, is usually referred to as a turbine stage and most known turbine engines include a plurality of turbine stages. The arrangement of turbine buckets and nozzle segments is referred to as a hot gas path.
Such known turbine buckets and nozzle segments in the hot gas path may wear over time. For example, such hot gas path components may exhibit stress-related cracking induced by temperatures at or above predetermined parameters. Therefore, many known gas turbine engines include temperature monitoring systems that provide operational temperature data in real time, i.e., at the time of measurement. At least some of these known temperature monitoring systems monitor and record temperature data as an input to adjust operation, e.g., the firing rate of the gas turbine engine, i.e., the rate and/or ratio of fuel and air being combusted in the engine. In some cases, the temperature data may be used as an input into certain protective features of the engine.
Measuring gas temperatures in a combusting flame or harsh environment downstream of a combustor, i.e., the hot gas path may include many sources of inaccuracy and non-repeatability. Many of those relate to physical properties of the temperature measurement mechanisms positioned in or proximate the flow of the hot combustion gases and/or proximate the high-temperature gas turbine components. For example, such detection mechanisms include thermocouples and gas sampling probes for point temperature measurements. However, these temperature measurement mechanisms do not account for radiation effects prominent in the hot gas path. Also, these temperature measurement mechanisms do not provide accurate temperature distribution profiles and alternative computational extrapolations and approximations must be used to facilitate spatial-resolution of the temperature profiles, albeit, with some inaccuracies induced by the modeling techniques and approximations used. At least some other known temperature measurement mechanisms include laser diagnostic techniques, e.g., laser Rayleigh scattering, laser Raman scattering, and planar laser induced fluorescence. However, these temperature measurement mechanisms are difficult to implement for temperature control of the gas turbine engine.
Therefore, to overcome the deficiencies of known temperature measurement mechanisms with respect to gas temperature profiles and near-wall temperature measurements in high-temperature and high-pressure environments, gas turbine manufacturers may elect to fabricate, install, and run hot gas components with greater thermal margins to extend the useful service life of such components. Increasing thermal margins typically manifests as increased wall thicknesses and other ruggedizing methods. Such increased ruggedness of those components increases the costs of production and increases a potential for premature reductions in service life due to excessive temperature profiles induced in the walls of the components during operations that typically include large-scale temperature changes, e.g., startups, shutdowns, and load changes. Increasing thermal margins during gas turbine operation is typically manifested as increased cooling flow rates. Increased cooling flow usage for those components increases the fuel consumption and decreases gas turbine efficiency.